1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a film cooling hole for a turbine airfoil.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
Airfoils used in a gas turbine engine, such as rotor blades and stator vanes (guide nozzles), require film cooling of the external surface where the hottest gas flow temperatures are found. The airfoil leading edge region is exposed to the highest gas flow temperature and therefore film cooling holes are used here. Film cooling holes discharge pressurized cooling air onto the airfoil surface as a layer that forms a blanket to protect the metal surface from the hot gas flow. The prior art is full of complex film hole shapes that are designed to maximize the film coverage on the airfoil surface while minimizing loses.
Film cooling holes with large length to diameter ratio are frequently used in the leading edge region to provide both internal convection cooling and external film cooling for the airfoil. For a laser or EDM formed cooling hole, the typical length to diameter is less than 12 and the film cooling hole angle is usually no less than 20 degrees relative to the airfoil's leading edge surface. FIGS. 1 and 2 show a prior art film cooling hole with a large length to diameter (L/D) ratio as discloses in U.S. Pat. No. 6,869,268 B2 issued to Liang on Mar. 22, 2005 and entitled COMBUSTION TURBINE WITH AIRFOIL HAVING ENHANCED LEADING EDGE DIFFUSION HOLES AND RELATED METHODS. In order to attain the same film hole breakout length or film coverage shown in FIG. 2, the straight circular showerhead hole in FIG. 1 has to be at around 14 degrees relative to the airfoil leading edge surface. This also results in a length to diameter ration of near 14. Both the film cooling hole angle and L/D exceed current manufacturing capability.
FIG. 2 shows a one dimension diffusion showerhead film cooling hole design which reduces the shallow angle required by the straight hole and changes the associated L/D ratio to a more producible level. This film cooling hole includes a constant diameter section at the entrance region of the hole that provides cooling flow metering capability, and a one dimension diffusion section with less than 10 degrees expansion in the airfoil radial inboard direction. As a result of this design, a large film cooling hole breakout is achieved and the airfoil leading edge film cooling coverage and film effectiveness level is increased over the FIG. 1 straight film cooling hole.
For an airfoil main body film cooling, a two dimensional compound shaped film hole as well as a two dimensional shaped film cooling hole with curved expansion is utilized to enhance film coverage and to minimize the radial over-expansion when these cooling holes are used in conjunction with a compound angle. U.S. Pat. No. 4,653,983 issued to Vehr on Mar. 31, 1987 and entitled CROSS-FLOW FILM COOLING PASSAGE and U.S. Pat. No. 5,382,133 issued to Moore et al on Jan. 17, 1995 and entitled HIGH COVERAGE SHAPED DIFFUSER FILM HOLE FOR THIN WALLS both disclose this type of film cooling hole.
A three dimensional diffusion hole in the axial or small compound angle and variety of expansion shape was also utilized in an airfoil cooling design for further enhancement of the film cooling capability. U.S. Pat. No. 4,684,323 issued to Field on Aug. 4, 1987 and entitled FILM COOLING PASSAGES WITH CURVED CORNERS and U.S. Pat. No. 6,183,199 B1 issued to Beeck et al on Feb. 6, 2001 and entitled COOLING-AIR BORE show this type of film hole.
Another improvement over the prior art three dimensional film hole is disclosed in U.S. Pat. No. 6,918,742 B2 issued to Liang on Jul. 19, 2005 and entitled COMBUSTION TURBINE WITH AIRFOIL HAVING MULTI-SECTION DIFFUSION COOLING HOLES AND METHODS OF MAKING SAME. This multiple diffusion compounded film cooling hole starts with a constant diameter cross section at the entrance region to provide for a cooling flow metering capability. The constant diameter metering section is followed by a 3 to 5 degree expansion in the radial outward direction and a combination of a 3 to 5 degree followed by a 10 degree multiple expansions in the downstream and radial inboard direction of the film hole. There is no expansion for the film hole on the upstream side wall where the film cooling hole is in contact with the hot gas flow.
U.S. Pat. No. 4,653,983 issued to Vehr on Mar. 31, 1987 and entitled CROSS FLOW FILM COOLING PASSAGE discloses a regular shaped film cooling hole of the prior art with the film ejection stream located above the airfoil surface in which vortices form underneath the film cooling discharge from the hole. The film cooling hole is the standard 10-10-10 expansion file hole where the two sides and the bottom of downstream side of the hole all have degrees of expansion. The film flow will penetrate into the main stream and then reattach to the airfoil surface at a distance of around 2 times the film hole diameter. Thus, hot gas injection into the space below the film injection location and subsequently a pair of vortices is formed under the film flow. As a result of the shear mixing, the film effectiveness is reduced. The film layer of cooling air reattaches to the airfoil surface downstream from the vortices that are formed.